In the operation of a gas turbine engine, air at atmospheric pressure is initially compressed by a compressor, and the resulting compressed air is delivered to a combustion stage. In the combustion stage, heat is added to the compressed air leaving the compressor by mixing fuel with the compressed air and by burning the fuel/air mixture. The gas flow resulting from the combustion of the fuel/air mixture in the combustion stage expands through a turbine, and some of the energy of the gas flow is used to drive a turbine in order to produce mechanical power.
One form of turbine is an axial turbine having one or more stages, wherein each stage employs one row of stationary nozzle guide vanes and one row of moving blades. The row of moving blades is mounted on a turbine disk. The nozzle guide vanes are aerodynamically designed to direct incoming gas from the combustion stage onto the turbine blades to thereby aerodynamically transfer kinetic energy to the blades.
In the past, the combustion gases entering the turbine typically have had a gas entry temperature in the range of 850.degree. to at least 1200.degree. F. Since the efficiency and work output of the turbine engine are related to the gas entry temperature of the incoming combustion gases, there is a trend in gas turbine engine technology to increase the gas entry temperature. A consequence of increasing the gas entry temperature of the combustion gases in a gas turbine engine is that the materials of the nozzle guide vanes and blades must be chosen so that the nozzle guide vanes and blades can resist such increased gas entry temperatures.
Historically, nozzle guide vanes and blades have been made of metals such as high temperature steels and, more recently, nickel alloys. Even with these types of high temperature materials, it has been found necessary to provide internal cooling passages in order to prevent melting of these materials. Also, ceramic coatings can be applied to the nozzle guide vanes and blades to enhance the heat resistance of such nozzle guide vanes and blades. In specialized applications, nozzle guide vanes and blades are being made entirely of ceramic, which resists even higher gas entry temperatures.
However, if the nozzle guide vanes and/or blades are made of ceramic, which has a different chemical composition, physical property, and coefficient of thermal expansion to that of a metal supporting structure such as the disk to which the blades are typically mounted, then undesirable stresses, a portion of which are thermal stresses, will result between the nozzle guide vanes and/or blades and their supports when the turbine engine is operating. Such undesirable thermal stresses cannot effectively be contained by cooling.
Furthermore, conventional joints between blades and disks of a turbine have typically used a fir tree, or a dove tail, root design. Historically, a dove tail root design has been used with a ceramic blade to attach the blade to a metallic disk. A metallic compliant layer of material is used between the highly stressed ceramic blade root and the metallic disk to accommodate the relative movement, and resulting sliding friction, that may occur. The sliding friction between the ceramic blade and the metallic disk creates a contact tensile stress on the ceramic that degrades the surface of the ceramic. This degradation in the surface of the ceramic occurs in a tensile stress zone of the blade root. Therefore, when a surface flaw is generated in the ceramic of critical size, the blade root fails catastrophically.
Other turbine wheel assemblies involve bonding a ceramic ring, having a plurality of ceramic blades integrally formed thereon, to a fully dense ceramic disk, or involve the use of a ceramic disk having ceramic blades integrally formed therewith. All such prior art arrangements have been developed based on the assumption that the disk material must react to both the centrifugal and aerodynamic loads which are imposed by and on the blades under normal operating conditions of the gas turbine engine. That is, since the blades and disk of these prior art turbine wheels are, for the most part, rigid structures, the aerodynamic loads imparted onto the blades by the expansion of the hot gases produced by the combustion stage, and the centrifugal forces produced by the blades from the resulting disk rotation, are transferred from the blades to the disk. While disk materials have primarily been either of metallic composition for engines with turbine inlet temperatures less than approximately 2200.degree. F., and have been monolithic or composite ceramic materials for turbine inlet temperatures greater than 2200.degree. F., the prior art turbine wheels which are designed on the assumption that the disk must react to both the centrifugal and aerodynamic forces of the blades have been massive structures.
The present invention overcomes one or more of the problems as set forth above.